Gas turbine engine



y 1968 e. L. WILDE ETAL 3,385,064

GAS TURBINE ENGINE Filed Jan. 5, 1967 I Inventor %;5%;u

y I M, 3 Attorneys United States Patent 3,385,664 GAS TURBINE ENGINE Geoffrey Light Wilde and James Alexander Petrie, Derby,

England, assignors to Rolls-Royce Limited, Derby,

England, a British company Filed Jan. 5, 1967, Ser. No. 667,451 Claims priority, application Great Britain, Jan. 7, 1366, 747/66 12 Claims. (Cl. 6il--226) ABSTRACT OF THE DTSCLQSURE A gas turbine engine having low and high pressure compressors connected by respective shafts to low and high pressure turbines, all housed within an inner casing, and a fan disposed outwardly of the inner casing and driven by an intermediate pressure turbine located between the high and low pressure turbines, and preferably also driving an intermediate pressure compressor.

This invention relates to gas turbine engines.

According to the present invention a gas turbine engine comprises an inner casing within which are disposed in flow series a low pressure compressor, a high pressure compressor, combustion equipment and high and low pressure turbine drivingly connected to the high and low pressure compressors respectively, and further comprising a fan disposed outwardly of the inner casing and drivingly connected to an intermediate pressure turbine which is independent of and intermediate the high and low pressure turbines.

Preferably, the fan is disposed axially between the high and low pressure compressors.

Preferably, an intermediate pressure compressor is disposed between the low and high pressure compressors and is drivingly connected to the fan and to the said intermediate pressure turbine.

Said fan may comprise one or more stages of blades which are connected to the radially outer ends of one or more rotor blades of the intermediate pressure compressor.

The fan blades may be disposed within an annular fan duct surrounding at least the intermediate pressure compressor.

The fan duct may have an annular air intake at its upstream end surrounding the air intake to the low pressure compressor. Said fan intake according to one embodiment of the invention is disposed downstream of the intake to the low pressure compressor.

Sealing means may be provided where the said fan blades extend through the inner casing substantially to preclude leakage of air from the intermediate pressure compressor through the inner casing.

The invention will now be described, by way of example only, with reference to the accompanying drawing which shows a part-section taken axially through an engine in accordance with the present invention.

The engine comprises an inner casing made up of front and rear casing parts 10 and 110 within which are disposed in flow series a low pressure compressor 11, an intermediate pressure compressor 12, a high pressure compressor 13, combustion equipment 14, a high pressure turbine 15, an intermediate pressure turbine 16 and a low pressure turbine 17. The low pressure compressor 11 and the low pressure turbine 17 are interconnected by a shaft 18 which is carried by a tail bearing 19 and a front bearing 20. The bearing 19 is supported from the casing part 110 by struts 22 and the front bearing is supported from the casing part 10 by struts 23.

The low pressure compressor comprises a supporting 3,335,@54 ?atenteel May 28, 1968 drum 24 which carries a number of rows of rotor blades 25 interposed between successive rows of stator blades 25, the stator blades 26 being carried by the casing part 10 and extending inwardly therefrom.

The low pressure turbine 17 comprises a turbine disc 27 carrying a single row of turbine blades 28. A row of guide vanes 29 is situated upstream of the turbine blades 28 and serves to direct gases onto the turbine blades 28.

The intermediate pressure compressor 12 and intermediate pressure turbine 16 are interconnected by an intermediate shaft 3%. The intermediate shaft 30 is supported by a tail bearing 31 which is carried by the shaft 18 and a front bearing 33 which is carried by struts 34 from the casing part 10. The compressor 12 comprises two rows of rotating blades 38 which are supported from the shaft 30 as described below and which are interposed between two rows of stator blades 39 carried on the outer surface of a drum 40 which is supported from the casing part 110 by struts 41 which also form part of a fixed structure 36. Between the outlet from the low pressure compressor 11 and the inlet to the intermediate pressure compressor 12 there is interposed a row of guide vanes 42 which are formed as continuations of respective struts 34.

The upstream row of rotor blades 38 is attached to the intermediate shaft 30 by a disc 37, each blade 38 carrying at its outer end a root 43 of a first row of fan blades 44. An axially extending cylindrical connecting member 45 is attached to the roots 43 and extends in a downstream direction. The member 45 supports at its downstream end the roots 46 of a second row of fan blades 47. From these roots 46 there extends inwardly the downstream row of rotor blades 38 of the intermediate pressure compressor 12. Thus the upstream row of rotor blades 38 carries the first row of fan blades 44 and, by way of the connecting member 45, the second row of fan blades 47 and the downstream rotor blades 38.

In order to prevent gases leaking from the working duct of the intermediate prmsure compressor into the fan, seals are provided between the rotating parts of the compressor and the casing parts 10 and 110. From the forward faces of the roots 43 there extends a first sealing member 48 which forms a seal with the casing part 10 and from the rearward faces of the roots 46 there extends a second sealing member 49 which forms a seal with the casing part 116. As the connecting member 45 is solid and is sealed to the roots 43 and 46 the working duct of the intermediate pressure compressor is thereby effectively sealed from the working duct of the fan.

The high pressure compressor 13 and the high pressure turbine 15 are interconnected by a high pressure shaft 50. The high pressure shaft 50 is supported from the intermediate shaft 30 by a tail bearing 51 and from the fixed structure 36 by a front bearing 52. The high pressure compressor 13 comprises a drum 53 which supports a number of rows of rotor blades 54 which are interposed between successive rows of stator blades 55 supported by the easing part 119 and extending inwardly therefrom. Between the intermediate pressure compressor 12 and the low pressure compressor 13 there are disposed guide vanes 56 and 57, guide vanes 56 forming parts of respective struts 41. At the outlet end of the high pressure compressor 13 is a row of guide vanes 58 at the entrance to the combustion equipment 14.

The combustion equipment 14 comprises inner and outer casings 59 and 60 defining between them an annu lar combustion chamber within which a flame tube 61 is mounted. At the upstream end of the flame tube 61 there is provided a number of fuel injectors 62 by which fuel can be injected into the air stream flowing through the fiame tube 61.

Surrounding the inner casing parts 10 and and extending coaxially therewith is an outer casing 65. This casing 65 is supported from the casing part 110 by a series of struts 68 which may form the outlet guide vanes of the fan, and by further supports 71 at the downstream end of the casing part 110. Between the two rows of fan blades 44 and 47 there is disposed a row of stator blades 69 which terminate at their innermost extremities in a shroud ring 70, the shroud ring forming a continuous extension of the outer surface of easing parts 10 and 116. The radially inwardly facing surface of the shroud ring 70 co-operates with a series of fins (not shown) on the outer surface of the connecting members 45 so as to provide an annularly extending seal and to prevent air from the outlet of the first row of fan blades 44 leaking to the inlet of the second row of fan blades 47 without being deflected by the stator blades 69.

The casing part 10 at its upstream end forms with a fixed nose cone 72 supported by the struts 23 an intake for the low pressure compressor 11. Similarly the upstream end of the outer casing 65 forms with the casing part 10 an annular intake for the fan.

In the illustrated embodiment the upstream end of the outer casing 65 lies some distance downstream of the upstream end of the casing part 16. In this way, since the casing part 10 is larger in cross-sectional area at its upstream end than it is further downstream, a given size of intake area for the fan can be achieved with the minimum diameter of the outer casing 65. Alternatively, the upstream end of the casing 65 may be disposed forwardly of the upstream end of the casing part 10 for other reasons such as noise suppression.

The inner casing downstream part 110 terminates downstream of the struts 22 in an inner propulsion nozzle 73 while the outer casing 65 terminates in an outer nozzle 74 which is either downstream (as shown) or upstream of the inner propulsion nozzle 73 according to the requirements of the engine installation.

In operation of the engine the low pressure, intermediate pressure and high pressure compressors in succession compress air which has been drawn into the intake of the low pressure compressor 11. The compressed air is then delivered to the combustion equipment 14 in which fuel is burnt and the resulting hot exhaust gases pass in turn through the high pressure, intermediate pressure and low pressure turbines 15, 16 and 17 to drive these turbines. The exhaust gases from the low pressure turbine 17 pass through the inner propulsion nozzle 73 to produce a forward propulsive thrust.

The rotor blades 38 of the intermediate pressure compressor 12, being attached to the fan blades 44 and 47, cause these blades to rotate, compressing air in the annular section duct formed between the inner casing 10, 110 and the outer casing 65. This air exhausts through the outer nozzle 74, augmenting the forward propulsive thrust produced by the engine.

The engine according to the invention has particular advantages when compared for instance with conventional front fan engines. Thus in a front fan engine having fan blades which are extensions of low pressure compressor blades, the air inlet is split into two concentric annuli, with the low speed root parts of the fan blades surrounding a relatively large hub which tends to block the intake. In the present invention the inner intake annulus is occupied by efficient blades surrounding a very small hub. The intermediate pressure compressor blades are supercharged by the low pressure compressor and therefore in comparison with the above-mentioned front fan engine the hub section of the fan can be of reduced cross-sectional area. The fan blade root speed is therefore higher and there is more radial depth available for rotor disc area radially inwardly of the fan blades to support the fan blade rotor system.

The present invention also helps to ease fan icing problems which would otherwise arise with the front fan engine since the intermediate pressure compressor blades 4 and hence the roots 43, 46 of the fan blades 44, 47 are heated by the compressed air from the low pressure compressor 11.

The fan blades 44, 47 are shorter and lighter compared with fan blades of conventional fan engines because of the larger ratio of the hub to tip diameters in the annular fan duct.

Again in the engine according to the present invention the fan is driven through an intermediate pressure shaft. This is evidently better than driving through the low pressure shaft since the latter must of necessity be of smaller diameter and therefore inherently less able to resist high torque than an intermediate pressure shaft. This should result in smaller bearings and a general savings in overall weight. Moreover, the fan is driven by the intermediate pressure turbine which is the most effective section in which the average hub/tip ratio is higher than that in a conventional low pressure turbine. Since the low pressure turbine drives only the low pressure compressor in the engine of the present invention it can be a single stage turbine which can have a relatively high rotational speed and is therefore more efiicient. The low pressure turbine is the coolest stage in the turbine, so that the relatively large low pressure turbine blades are able to withstand the high stresses resulting from their high rotational speed.

As a modification to the turbine means illustrated a row of variable guide vanes could be interposed between the intermediate pressure turbine 16 and the low pressure turbine 17. This would give the advantage that the low pressure/ intermediate pressure turbine speed ratio could be varied. In this way overall pressure ratio of the engine could be maintained with some variation of relative mass flow through the fan, since an increase in low pressure nozzle guide vane angle would tend to decrease fan speed and increase the low pressure compressor speed. This feature might be of great value in controlling noise emission from the fan under certain flight conditions, such as approach to land and landing.

As described above with reference to the drawing the air inlet for the fan can be tucked in behind the inlet for the low pressure compressor to a certain extent, in this way reducing the external diameter of the engine pod as much as possible. This, combined with the reduction in size of the low pressure compressor and its hub, enables the engine to be made small for a given mass flow. Furthermore, the very low speed parts of blading of a conventional high by-pass engine, that is, the root parts of the front fan and the low pressure turbine, are eliminated. These are now replaced by blade sections running at optimum speeds, which are therefore smaller and lighter for their duty.

In the proposal described adjacent shafts are contrarotational relative to each other. This feature could be exploited in the turbine by enabling guide vanes between adjacent rotor stages to be dispensed with.

We claim:

1. A gas turbine engine comprising an inner casing, a low pressure compressor, a high pressure compressor, combustion equipment, a high pressure turbine, an intermediate pressure turbine and a low pressure turbine disposed in axial flow series in said inner casing, the high and low pressure turbines being drivingly connected by concentric shafts to the high and low pressure compressors respectively, and further comprising a fan which is disposed outwardly of the inner casing and which is drivingly connected to the intermediate pressure turbine.

2. A gas turbine engine as claimed in claim 1 in which the fan is disposed axially between the high and low pressure compressors.

3. A gas turbine engine as claimed in claim 2 in which an intermediate pressure compressor is disposed between the low and high pressure compressors and is drivingly connected to the fan and to the said intermediate pressure turbine.

4. A gas turbine engine according to claim 3 in which the fan comprises at least one stage of blades connected to the radially outer ends of at least one rotor blade of the intermediate pressure compressor.

5. A gas turbine engine as claimed in claim 3 in which an annular fan duct is provided and surrounds at least the intermediate pressure compressor, the fan being disposed within said fan duct.

6. A gas turbine engine according to claim 5 in which the fan duct and the low pressure compressor have respective air intakes at their upstream ends, the air intake to the fan surrounding the air intake to the low pressure compressor.

7. A gas turbine engine according to claim 6 in which the fan intake is disposed downstream of the intake to the low pressure compressor.

8. A gas turbine engine as claimed in claim 4 in which sealing means are provided where the said fan blades extend through the inner casing substantially to preclude leakage of air from the intermediate pressure compressor through the inner casing.

9. A gas turbine engine as claimed in claim 3 in which a shaft is drivingly connected to the intermediate pressure turbine and the intermediate pressure compressor comprises a rotor disc mounted on said shaft, a plurality of stages of rotor blades, and a connecting member, one of the stages of rotor blades being attached to the rotor disc and the remaining stage(s) being drivingly connected to the said one stage by the connecting member, the connecting member interconnecting the radially outer ends of the blades of said stages.

10. A gas turbine engine according to claim 9 in which the fan has a plurality of stages of blades one of which is directly connected to the radially outer ends of the blades of said one stage of the intermediate pressure compressor and the remainder of which are connected to the said connecting member.

11. A gas turbine engine as claimed in claim 1 in which a row of variable incidence guide vanes is mounted upstream of the low pressure turbine.

12. A gas turbine engine as claimed in claim 3 in which respective nested shafts carry the low pressure compressor and turbine, the intermediate pressure compressor and turbine and the high pressure compressor and turbine, adjacent ones of said shafts being contra-rotational.

References Cited UNITED STATES PATENTS 11 1947 Heppner 230122 9/1966 Hull -226 

